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Abstract:

A method of calculating a throat area of a section of a machinery,
according to an exemplary aspect of the present disclosure includes,
among other things, outlining a boundary of the throat area of the
section, selecting a plurality of inspection points along the boundary of
the throat area, dividing the throat area into a plurality of sections,
calculating an individual area of each of the plurality of sections and
summing the individual areas of each of the plurality of sections to
calculate the throat area.

Claims:

1. A method associated for use with a gas turbine engine, comprising:
calculating a throat area of a vane segment of the gas turbine engine by
radially dividing the throat area into a plurality of sections,
calculating an individual area associated with each of the plurality of
sections, and summing the individual areas to calculate the throat area,
the vane segment including a first platform and a second platform and a
first airfoil and a second airfoil extending between the first platform
and the second platform; and adjusting a design characteristic of the
vane segment based on the calculated throat area.

2. The method as recited in claim 1, wherein adjusting the design
characteristic of the vane segment includes cutting or trimming a portion
of the vane segment.

3. The method as recited in claim 2, wherein cutting or trimming the
portion of the vane segment includes trimming a trailing edge of at least
one of the first airfoil and the second airfoil.

4. The method as recited in claim 2, wherein cutting or trimming the
portion of the vane segment includes machining the portion to remove
material.

5. The method as recited in claim 2, wherein cutting or trimming the
portion of the vane segment includes grinding the portion to remove
material.

6. The method as recited in claim 2, wherein cutting or trimming the
portion of the vane segment includes electrical discharge machining the
portion to remove material.

7. The method as recited in claim 1, comprising, prior to radially
dividing the throat area, outlining a boundary of the throat area of the
vane segment by referencing a CAD model of the vane segment.

8. The method as recited in claim 7, comprising selecting a plurality of
inspection points along the boundary of the throat area.

9. The method as recited in claim 8, comprising measuring a location of
each of the plurality of inspection points using a coordinate measuring
machine (CMM).

10. The method as recited in claim 1, comprising comparing the calculated
throat area to a desired throat area and adjusting the design
characteristic of the vane segment only if the calculated throat area is
outside a manufacturing tolerance related to the desired throat area.

11. A method associated for use with a gas turbine engine, comprising:
inspecting a vane segment of the gas turbine engine, the vane segment
including an end wall contoured flow path, and the vane segment including
a first platform and a second platform and a first airfoil and a second
airfoil extending between the first platform and the second platform;
outlining a boundary of a throat area of the vane segment by referencing
a CAD model of the vane segment; selecting a plurality of inspection
points along the boundary of the throat area, at least a portion of the
plurality of inspection points being located along the end wall contoured
flow path of the vane segment; measuring a location of each of the
plurality of inspection points on an actual casting of the vane segment
using a coordinate measuring machine (CMM); calculating the throat area
of the vane segment, including portions of the throat area encompassing
the end wall contoured flow path, by radially dividing the throat area
into a plurality of sections, calculating an individual area associated
with each of the plurality of sections, and summing the individual areas
to calculate the throat area; and adjusting a design characteristic of
the vane segment if the calculated throat area is different than a
desired throat area.

12. The method as recited in claim 11, wherein the end wall contoured
flow path includes a plurality of undulations.

13. The method as recited in claim 11, comprising adjusting the throat
area calculation based on an error value associated the calculated throat
area.

14. The method as recited in claim 11, wherein adjusting the design
characteristic of the vane segment includes cutting or trimming a
trailing edge of either the first airfoil or the second airfoil.

15. The method as recited in claim 14, wherein cutting or trimming the
trailing edge includes machining or grinding the trailing edge to remove
material from the vane segment.

Description:

CROSS-REFERENCE TO RELATED APPLICATIONS

[0001] This application is a continuation-in-part of U.S. patent
application Ser. No. 13/630,145, which was filed Sep. 28, 2012.

BACKGROUND

[0002] This disclosure relates to a gas turbine engine, and more
particularly to a method of calculating a throat area of a section of a
gas turbine engine.

[0003] Gas turbine engines typically include a compressor section, a
combustor section and a turbine section. In general, during operation,
air is pressurized in the compressor section and is mixed with fuel and
burned in the combustor section to generate hot combustion gases. The hot
combustion gases flow through the turbine section, which extracts energy
from the hot combustion gases to power the compressor section and other
gas turbine engine loads.

[0004] The compressor and turbine sections of the gas turbine engine
typically include alternating rows of rotating blades and stationary
vanes. The rotating blades either create or extract energy from the hot
combustion gases that are communicated through the gas turbine engine,
and the vanes convert the velocity of the airflow into pressure and
prepare the airflow for the next set of blades.

[0005] The vanes are typically arranged in an array such that adjacent
vanes of a vane assembly are circumferentially disposed about a
longitudinal centerline axis of the gas turbine engine. The hot
combustion gases that are communicated through the gas turbine engine
flow through a throat area that extends between the adjacent vanes of the
vane assembly. The throat area controls the pressure ratio and mass flow
rate of the gas turbine engine.

[0006] The throat area associated with a particular section of a gas
turbine engine, such as a turbine section, can be calculated using a
series of gauges that measure the location of a pressure side and suction
side of each airfoil of a vane assembly associated with the turbine
section. These measurements can be used to mathematically calculate a
throat area associated with the vane assembly. This method assumes
axi-symmetric flow path surfaces and ignores any undulations or flow path
end wall contouring of the platforms of the vane assembly, which
potentially invalidate the calculation.

SUMMARY

[0007] A method of calculating a throat area of a section of a machinery,
according to an exemplary aspect of the present disclosure includes,
among other things, outlining a boundary of the throat area of the
section, selecting a plurality of inspection points along the boundary of
the throat area, dividing the throat area into a plurality of sections,
calculating an individual area of each of the plurality of sections and
summing the individual areas of each of the plurality of sections to
calculate the throat area.

[0008] In a further non-limiting embodiment of the foregoing method of
calculating a throat area of a section of a machinery, the step of
outlining includes outlining the boundary of the throat area using a CAD
model of the section.

[0009] In a further non-limiting embodiment of either of the foregoing
methods of calculating a throat area of a section of a machinery, the
step of dividing includes radially dividing the throat area into the
plurality of sections.

[0010] In a further non-limiting embodiment of any of the foregoing
methods of calculating a throat area of a section of a machinery, the
plurality of sections are trapezoidal shaped.

[0011] In a further non-limiting embodiment of any of the foregoing
methods of calculating a throat area of a section of a machinery, the
plurality of sections are triangular shaped.

[0012] In a further non-limiting embodiment of any of the foregoing
methods of calculating a throat area of a section of a machinery, a first
portion of the plurality of sections are trapezoidal shaped and a second
portion of the plurality of sections are triangular shaped.

[0013] In a further non-limiting embodiment of any of the foregoing
methods of calculating a throat area of a section of a machinery, the
method includes calculating an error value is associated with the throat
area.

[0014] In a further non-limiting embodiment of any of the foregoing
methods of calculating a throat area of a section of a machinery, the
step of adjusting the throat area is based on the error value.

[0015] In a further non-limiting embodiment of any of the foregoing
methods of calculating a throat area of a section of a machinery, the
step of calculating includes measuring a distance between a first portion
of the plurality of inspections points located on an outer diameter
portion of the boundary and a second portion of the plurality of
inspection points located on an inner diameter portion of the boundary.

[0016] In a further non-limiting embodiment of any of the foregoing
methods of calculating a throat area of a section of a machinery, the
step of measuring a location of each of the plurality of inspection
points uses a coordinate measuring machine (CMM).

[0017] In a further non-limiting embodiment of any of the foregoing
methods of calculating a throat area of a section of a machinery, the
section is a mid-turbine frame of a gas turbine engine.

[0018] In a further non-limiting embodiment of any of the foregoing
methods of calculating a throat area of a section of a machinery, the
section is a turbine section of a gas turbine engine.

[0019] A method of calculating a throat area of a section of a gas turbine
engine, according to another exemplary aspect of the present disclosure
includes, among other things, radially dividing a throat area of the
section into a plurality of sections, calculating an individual area
associated with each of the plurality of sections of the throat area and
summing the individual areas of each of the plurality of sections to
calculate the throat area.

[0020] In a further non-limiting embodiment of the foregoing method of
calculating a throat area of a section of a gas turbine engine, the
section is a turbine section of the gas turbine engine.

[0021] In a further non-limiting embodiment of either the foregoing
methods of calculating a throat area of a section of a gas turbine
engine, the method includes outlining a boundary of the throat area of
the turbine section in a CAD model prior to the step of radially
dividing.

[0022] In a further non-limiting embodiment of any of the foregoing
methods of calculating a throat area of a section of a gas turbine
engine, the method includes selecting a plurality of inspection points
along the boundary of the throat area.

[0023] In a further non-limiting embodiment of any of the foregoing
methods of calculating a throat area of a section of a gas turbine
engine, the boundary includes an outer diameter portion, an inner
diameter portion, a suction side portion and a pressure side portion and
the plurality of inspection points are selected on each of the outer
diameter portion, the inner diameter portion, the suction side portion
and the pressure side portion.

[0024] In a further non-limiting embodiment of any of the foregoing
methods of calculating a throat area of a section of a gas turbine
engine, the step of radially dividing includes measuring a radial
distance between a plurality of outer diameter portions and a plurality
of inner diameter portions of a boundary of the throat area.

[0025] In a further non-limiting embodiment of any of the foregoing
methods of calculating a throat area of a section of a gas turbine
engine, the section includes an end wall having at least one undulation.

[0026] In a further non-limiting embodiment of any of the foregoing
methods of calculating a throat area of a section of a gas turbine
engine, the method includes adjusting the throat area calculation based
on an error value associated with a trailing edge location of the
section.

[0027] The various features and advantages of this disclosure will become
apparent to those skilled in the art from the following detailed
description. The drawings that accompany the detailed description can be
briefly described as follows.

[0029] FIGS. 2A and 2B illustrate a segment of a section of a gas turbine
engine.

[0030] FIG. 3 schematically illustrates a throat area associated with a
segment of a section of a gas turbine engine.

[0031] FIG. 4 schematically illustrates a method of calculating a throat
area of a segment of a section of a gas turbine engine.

DETAILED DESCRIPTION

[0032] FIG. 1 schematically illustrates a gas turbine engine 20. The
exemplary gas turbine engine 20 is a two-spool turbofan engine that
generally incorporates a fan section 22, a compressor section 24, a
combustor section 26 and a turbine section 28. Alternative engines might
include an augmentor section (not shown) among other systems for
features. The fan section 22 drives air along a bypass flow path B, while
the compressor section 24 drives air along a core flow path C for
compression and communication into the combustor section 26. The hot
combustion gases generated in the combustor section 26 are expanded
through the turbine section 28. Although depicted as a turbofan gas
turbine engine in the disclosed non-limiting embodiment, it should be
understood that the concepts described herein are not limited to turbofan
engines and these teachings could extend to other types of engines,
including but not limited to, three-spool engine architectures.

[0033] The gas turbine engine 20 generally includes a low speed spool 30
and a high speed spool 32 mounted for rotation about an engine centerline
longitudinal axis A. The low speed spool 30 and the high speed spool 32
may be mounted relative to an engine static structure 33 via several
bearing systems 31. It should be understood that other bearing systems 31
may alternatively or additionally be provided.

[0034] The low speed spool 30 generally includes an inner shaft 34 that
interconnects a fan 36, a low pressure compressor 38 and a low pressure
turbine 39. The inner shaft 34 can be connected to the fan 36 through a
geared architecture 45 to drive the fan 36 at a lower speed than the low
speed spool 30. The high speed spool 32 includes an outer shaft 35 that
interconnects a high pressure compressor 37 and a high pressure turbine
40. In this embodiment, the inner shaft 34 and the outer shaft 35 are
supported at various axial locations by bearing systems 31 positioned
within the engine static structure 33.

[0035] A combustor 42 is arranged between the high pressure compressor 37
and the high pressure turbine 40. A mid-turbine frame 44 may be arranged
generally between the high pressure turbine 40 and the low pressure
turbine 39. The mid-turbine frame 44 can support one or more bearing
systems 31 of the turbine section 28. The mid-turbine frame 44 may
include one or more airfoils 46 that extend within the core flow path C.

[0036] The inner shaft 34 and the outer shaft 35 are concentric and rotate
via the bearing systems 31 about the engine centerline longitudinal axis
A, which is co-linear with their longitudinal axes. The core airflow is
compressed by the low pressure compressor 38 and the high pressure
compressor 37, is mixed with fuel and burned in the combustor 42, and is
then expanded over the high pressure turbine 40 and the low pressure
turbine 39. The high pressure turbine 40 and the low pressure turbine 39
rotationally drive the respective high speed spool 32 and the low speed
spool 30 in response to the expansion.

[0037] In a non-limiting embodiment, the gas turbine engine 20 is a
high-bypass geared aircraft engine. In a further example, the gas turbine
engine 20 bypass ratio is greater than about six (6:1). The geared
architecture 45 can include an epicyclic gear train, such as a planetary
gear system or other gear system. The example epicyclic gear train has a
gear reduction ratio of greater than about 2.3, and in another example is
greater than about 2.5:1. The geared turbofan enables operation of the
low speed spool 30 at higher speeds, which can increase the operational
efficiency of the low pressure compressor 38 and low pressure turbine 39
and render increased pressure in a fewer number of stages.

[0038] The pressure ratio of the low pressure turbine 39 can be pressure
measured prior to the inlet of the low pressure turbine 39 as related to
the pressure at the outlet of the low pressure turbine 39 and prior to an
exhaust nozzle of the gas turbine engine 20. In one non-limiting
embodiment, the bypass ratio of the gas turbine engine 20 is greater than
about ten (10:1), the fan diameter is significantly larger than that of
the low pressure compressor 38, and the low pressure turbine 39 has a
pressure ratio that is greater than about five (5:1). It should be
understood, however, that the above parameters are only exemplary of one
embodiment of a geared architecture engine and that the present
disclosure is applicable to other gas turbine engines, including direct
drive turbofans.

[0039] In this embodiment of the exemplary gas turbine engine 20, a
significant amount of thrust is provided by the bypass flow path B due to
the high bypass ratio. The fan section 22 of the gas turbine engine 20 is
designed for a particular flight condition--typically cruise at about 0.8
Mach and about 35,000 feet. This flight condition, with the gas turbine
engine 20 at its best fuel consumption, is also known as bucket cruise
Thrust Specific Fuel Consumption (TSFC). TSFC is an industry standard
parameter of fuel consumption per unit of thrust.

[0040] Fan Pressure Ratio is the pressure ratio across a blade of the fan
section 22 without the use of a Fan Exit Guide Vane system. The low Fan
Pressure Ratio according to one non-limiting embodiment of the example
gas turbine engine 20 is less than 1.45. Low Corrected Fan Tip Speed is
the actual fan tip speed divided by an industry standard temperature
correction of [(Tram °R)/(518.7°R)]0.5, where T
represents the ambient temperature in degrees Rankine. The Low Corrected
Fan Tip Speed according to one non-limiting embodiment of the example gas
turbine engine 20 is less than about 1150 fps (351 m/s).

[0041] Each of the compressor section 24 and the turbine section 28 may
include alternating rows of rotor assemblies and vane assemblies (shown
schematically) that carry airfoils that extend into the core flow path C.
For example, the rotor assemblies can carry a plurality of rotating
blades 25, while each vane assembly can carry a plurality of vanes 27
that extend into the core flow path C. The blades 25 of the rotor
assemblies create or extract energy (in the form of pressure) from the
core airflow that is communicated through the gas turbine engine 20 along
the core flow path C. The vanes 27 of the vane assemblies direct the core
airflow to the blades 25 to either add or extract energy.

[0042] FIGS. 2A and 2B illustrate a vane segment 50 that can be
incorporated into a section 100 of a gas turbine engine, such as the gas
turbine engine 20 of FIG. 1. In this embodiment, the section 100 is the
turbine section 28 of the gas turbine engine 20 and the vane segment 50
is part of a vane assembly disposed within the turbine section 28.
However, other sections of a gas turbine engine 20 could also benefit
from the teachings of this disclosure, including but not limited to, the
compressor section and the mid-turbine frame.

[0043] The vane segment 50 can include a single annular structure or a
plurality of vane segments 50 could be mechanically attached to one
another and annularly disposed about the engine centerline axis A to form
a full-ring vane assembly. The vane segment 50 could include one or more
airfoils (or vanes) 56. In this particular embodiment, the section 100
includes a first vane segment 50A and a second vane segment 50B each
having a single airfoil 56.

[0044] The first and second vane segments 50A, 50B include an outer
diameter platform 52 and an inner diameter platform 54 spaced from the
outer diameter platform 52. The airfoils 56 extend between the outer
diameter platform 52 and the inner diameter platform 54. The airfoils 56
may extend radially across an annulus between the outer diameter platform
52 and the inner diameter platform 54. Each airfoil 56 includes a leading
edge 58, a trailing edge 60, a pressure side 62 and a suction side 64.
The airfoils 56 axially extend between the leading edge 58 and the
trailing edge 60 and circumferentially extend between the pressure side
62 and the suction side 64.

[0045] A gas path 65 is communicated axially downstream through the gas
turbine engine 20 along the core flow path C (FIG. 1) in a direction that
extends from the leading edges 58 toward the trailing edges 60 of the
airfoils 56. The gas path 65 represents the communication of core airflow
across the airfoils 56 and through the spaces that extend between the
adjacent airfoils 56 of the first and second vane segments 50A, 50B. The
space that extends between adjacent airfoils 56 of the vane segments 50A,
50B is referred to as a throat area 66.

[0046] As best illustrated in FIG. 2B, the inner diameter platform 54 and
the outer diameter platform 52 of one or both of the first and second
vane segments 50A, 50B include end walls 67 that establish the radially
outer and radially inner flow path boundaries of the gas path 65. One or
more of the end walls 67 can include undulations 69 that provide a
contoured end wall surface. The undulations 69 may reduce friction losses
that can occur as the airflow of the gas path 65 is communication through
the section 100, thereby increasing engine efficiency. However, these
undulations 69 can also add difficulty in calculating the throat area 66
using known throat area calculation methodologies.

[0047] FIG. 3 schematically illustrates a throat area 66 of a vane segment
50. The throat area 66 extends between a first airfoil 56A and a second,
adjacent airfoil 56B of the vane segment 50. The throat area 66 of the
vane segment 50 is a planar opening with a periphery bounded in a radial
direction by the outer diameter platform 52 (shown in FIG. 2) and the
inner diameter platform 54 and peripherally bounded in the
circumferential direction by the suction side 64 of the first airfoil 56A
and the trailing edge 60 of the second airfoil 56B (at the pressure side
62). Of course, a full ring vane assembly would include a plurality of
throat areas that extend between adjacent airfoils. The throat area 66 of
the vane segment 50 is an important parameter that may influence engine
efficiency. Therefore, the throat area 66 must be periodically measured
and calculated to verify clearances in the core flow path C.

[0048] FIG. 4, with continued reference to the previous Figures,
schematically illustrates an exemplary method 101 of calculating a throat
area 66 of a section of a gas turbine engine. In one embodiment, the
method 101 can be used to calculate a throat area 66 associated with a
turbine vane assembly of the gas turbine engine, although other sections
could also benefit from the teachings of this disclosure. The method 101
may be particularly useful for calculating a throat area on segments of
the gas turbine engine that include flow path end wall contouring (see,
e.g., undulations 69 of end walls 67 illustrated by FIG. 2B).

[0049] The method may begin at step block 102 by outlining a boundary 68
of the throat area 66. In one embodiment, the boundary 68 is outlined
using a CAD model of a section of the gas turbine engine. A person of
ordinary skill in the art having the benefit of this disclosure would be
able to select and outline the boundary 68 of the throat area 66 by
referencing a CAD model.

[0050] The boundary 68 is an outer periphery of the throat area 66. In
this embodiment, the boundary 68 extends radially between an outer
diameter portion 70 and an inner diameter portion 72 and extends
circumferentially between a suction side portion 74 and a pressure side
portion 76. The outer diameter portion 70 and the inner diameter portion
72 represent the locations where the throat area 66 is bounded by the
outer diameter platform 52 and the inner diameter platform 54,
respectively (see FIGS. 2A, 2B). The outer diameter portion 70 and the
inner diameter portion 72 may be contoured to reflect any undulations 69
formed in the end walls 67 of the outer diameter platform 52 and/or the
inner diameter platform 54. In other words, the outer diameter portion 70
and the inner diameter portion 72 may extend non-linearly as shown. In
addition, the suction side portion 74 of the boundary 68 represents the
portion of the throat area 66 that extends adjacent a suction side 64 of
a first airfoil 56A of the section and the pressure side portion 76
represents the portion of the throat area 66 that extends adjacent the
pressure side 62 of a second, adjacent airfoil 56B of the section (see,
for example, the vane segment 50 of FIG. 3). The suction side portion 74
and the pressure side portion 76 are linear in this embodiment.

[0051] Next, at step block 104, a plurality of inspection points (marked
in FIG. 4 with a + sign) are selected along the boundary 68 of the throat
area 66. For example, a plurality of inspection points OD1 through
ODN are selected along the outer diameter portion 70, a plurality of
inspection points ID1 through IDN are selected along the inner
diameter portion 72, a plurality of inspection points SS1 through
SSN are selected along the suction side portion 74 and a plurality
of inspection points PS1 through PSN are selected along the
pressure side portion 76. In one embodiment, the locations of these
various inspection points may be specified in an engineering print or CAD
model of the section of the gas turbine engine. However, the plurality of
inspection points may also be selected using other known methodologies.

[0052] The plurality of inspection points of the boundary 68 of the throat
area 66 can be measured at step block 106. For example, each of the
plurality of inspection points can be measured on an actual casting of
the section of the gas turbine engine using a coordinate measuring
machine (CMM). Other measuring techniques are also contemplated as within
the scope of this disclosure. The actual locations of the plurality of
inspection points of the boundary 68 can be recorded at step block 108,
including, to the extent the location is incorrect, a recordation of an
out-of-nominal distance associated with each of the plurality of
inspection points. As discussed below, the CMM data is later used to
calculate the area of a plurality of sections of the throat area 66.

[0053] At step block 110, the throat area 66 is divided into a plurality
of sections S1 through SN. In one embodiment, the throat area
66 is radially divided into the plurality of sections SN by drawing
radial lines L1 through LN from at least a first portion of the
plurality of inspection points OD1 through ODN to a
corresponding second portion of the plurality of inspection points
ID1 through ODN. In this embodiment, the inspection points
OD2 and ID2 are connected by radial line L2 that defines a
radial distance, inspection points OD4 and ID4 are connected by
radial line L4 that defines a radial distance, inspection points
OD6 and ID6 are connected by radial line L6 that defines a
radial distance, inspection points OD10 and ID10 are connected
by radial line L10 that defines a radial distance, inspection points
OD12 and ID12 are connected by radial line L12 that
defines a radial distance, and inspection points OD13 and ID13
are connected by a radial line L13 that defines a radial distance.
This is but one example of how the throat area 66 could be divided, and
it should be understood that the throat area 66 could be divided into
greater or fewer sections than are illustrated in this embodiment. The
division of the throat area 66 into the plurality of sections S1
through SN may also depend on the design of the section for which
the throat area is being calculated. That is, the throat areas 66 of
different parts may be divided into different numbered and shaped
sections.

[0054] The plurality of sections S1 through SN that divide the
throat area 66 may embody a variety of shapes. For example, the plurality
of sections S1 through SN can be trapezoidal shaped. In another
embodiment, the plurality of sections S1 through SN can be
triangular shaped. In yet another embodiment, a first portion of the
plurality of sections S1 through SN are trapezoidal shaped and
a second portion of the plurality of sections S1 through SN are
triangular shaped sections. Other shapes are also contemplated. For
example, each shape may be one in which an area of the shape can be
easily mathematically calculated.

[0055] Next, at step block 112, an individual area of each of the
plurality of sections S1 through SN is calculated. A person of
ordinary skill in the art having the benefit of this disclosure would
understand how to calculate the area of each of the plurality of sections
S1 through SN. For example, as is known, the area of a triangle
can be calculated using the formula A=1/2(b)(h), where b is the length of
the base of the triangle and h is the height of the triangle. In another
example, the area of a trapezoid can be calculated using the formula
1/2(b1+b2)h, where b1 and b2 are the lengths of the
bases and h is the height. In this embodiment, the lengths of the radial
lines L1 through LN can be used as the "h" values in the area
formulas. These radial distances are measured and calculated from the CMM
inspection data from step block 108. At step block 114, the individual
areas of each of the plurality of sections S through SN can be
summed to calculate the throat area 66.

[0056] An error value associated with the throat area 66 calculation can
optionally be calculated at step block 116. It may be necessary to
calculate the error value associated with the throat area 66 calculation
because the airfoil 56 of a section can be cast with a trailing edge 60
that is shorter or longer (due to shrinkage error) than defined by the
engineering print or CAD model of the section. The calculated throat area
66 can be adjusted based on the error value calculated at step block 116.
The error value may also be added to account for slight inabilities of
the formula to measure the exact flow area value that can result from a
coarse number of divisions of the flow area surface. Additional methods
may be incorporated to account for throat area 66 value effects from the
trailing edge 60 location.

[0057] At step block 118, the calculated throat area is compared to a
desired throat area. The desired throat area is a predefined value and
may be based on a desired pressure ratio and mass flow rate of the gas
turbine engine section, among other design criteria. In another
non-limiting embodiment, the desired throat area is established by aero
performance engineers in order to optimize the efficiency of the gas
turbine engine. The throat area impacts fuel efficiency, cost to operate
and durability of the gas turbine engine.

[0058] If the calculated throat area is outside of an acceptable tolerance
compared to the desired throat area, a design characteristic of the
section of the gas turbine engine can be modified at step block 120. For
example, adjusting the design characteristic of the gas turbine engine
section may include cutting or trimming a portion of the section to
remove material from the section. In one non-limiting embodiment, the gas
turbine engine section is a vane segment and cutting or trimming includes
trimming a trailing edge of an airfoil of the vane segment. The cutting
or trimming procedure may be accomplished by machining, grinding, or
electrical discharge machining the section to remove material from the
section and thereby influence the throat area 66 of the section. The
procedure described with reference to FIG. 4 may be particularly useful
for confirming that parts being received from vendors are within
specified manufacturing tolerances.

[0059] Although the different non-limiting embodiments are illustrated as
having specific components, the embodiments of this disclosure are not
limited to those particular combinations. It is possible to use some of
the components or features from any of the non-limiting embodiments in
combination with features or components from any of the other
non-limiting embodiments.

[0060] It should be understood that like reference numerals identify
corresponding or similar elements throughout the several drawings. It
should also be understood that although a particular component
arrangement is disclosed and illustrated in these exemplary embodiments,
other arrangements could also benefit from the teachings of this
disclosure.

[0061] The foregoing description shall be interpreted as illustrative and
not in any limiting sense. A worker of ordinary skill in the art would
understand that certain modifications could come within the scope of this
disclosure. For these reasons, the following claims should be studied to
determine the true scope and content of this disclosure.